1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to an air cooled blade outer air seal (BOAS) with cooled grooves for an industrial gas turbine engine.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A row or stage of turbine rotor blades rotate within an annular arrangement of ring segments in which blade tips form a small gap with an inner or hot surface of each ring segment. The size of the gap changes due to different thermal properties of the blade and the BOAS or ring segments from a cold state to a hot state of the turbine. The smaller the gap, the less hot gas leakage will flow between the blade tips and the ring segments.
An IGT engine operates for long periods of time at steady state conditions, as opposed to an aero gas turbine engine that operates for only a few hours before shutting down. Thus, the parts in the IGT engine must be designed for normal operation for these long periods, such as up to 40,000 hours of operation at steady state conditions.
High temperature turbine blade tip shroud heat load is a function of the blade tip section leakage flow. A high leakage flow will induce a high heat load on the blade tip shroud. Therefore, blade tip shroud cooling and sealing issues must be considered as a single problem. Prior art grooved turbine blade tip shroud includes a number of grooves opening from an underside surface of the tip shroud at from 90 to 130 degrees angle relative to the tip shroud backing structure in which the grooves extends into the flow path for the entire axial length of the blade outer air seal. The main purpose for incorporating grooved tip shroud in a blade design is to reduce the blade tip leakage and to provide for rubbing capability fro the blade tips. Prior art grooved blade tip shrouds used in the turbine design form straight teeth and are not cooled.
FIG. 1 shows a prior art turbine with a tip shroud cooling design. The blade tip shroud 11 includes film cooling holes that discharge cooling air from an impingement cavity 13 located between the tip shroud 11 and an impingement plate 14 that has impingement 15 holes spaced about it. During engine operation, the film cooling holes will be smeared when the blade tips rub into the tip shroud. Smearing of the film holes results in plugging of significantly reducing the cooling air flow to the tip shroud, and therefore reduces the cooling effectiveness of the tip shroud. An over-metal temperature will occur and the hot spots will cause erosion and shorten the part life of the engine. This is especially a problem in aero engines because the blade tip shroud is relatively small (when compared to an IGT engine) so that the film cooling holes are very small.